Outer diameter platform cooling hole system and assembly

ABSTRACT

An airfoil component is described herein. The airfoil component may include an OD platform comprising a gaspath face and a non-gaspath face coupled together via a plurality of cooling holes. The airfoil component may include an airfoil extending in from the outer diameter platform. The plurality of cooling holes comprises a plurality of groups of cooling holes disposed in the outer diameter platform proximate a suction side of the airfoil and a pressure side of the airfoil.

FIELD

This disclosure relates to a gas turbine engine, and more particularlyto cooling hole arrangements on a first stage vane outer diameter (OD)platform.

BACKGROUND

Gas turbine engines typically include a compressor section, a combustorsection and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

A typical turbine section includes at least one array of stator vanesarranged circumferentially about an engine central longitudinal axis todefine an outer radial flow path boundary for the hot combustion gasesleaving the combustor section and entering the turbine section. Thus,the convex outer diameter of the vane is an area where thecircumferential hotspot gas temperature is observed and is directlylinked to thermal distress due to the hot combustion gases.

SUMMARY

According to various embodiments, an airfoil component is describedherein. The airfoil component may include a platform, such as an outerdiameter platform, comprising a gaspath and a non-gaspath face/surface.The airfoil component may include an airfoil extending from the ODplatform. The plurality of cooling holes may comprise a plurality ofgroups of cooling holes disposed in the outer diameter platformproximate a suction side of the airfoil and a pressure side of theairfoil. A plenum may be formed adjacent to the non-gaspath face. Theplurality of cooling holes may be coupled to the plenum.

According to various embodiments, the airfoil component may be a firststage turbine vane. A first group of the plurality of cooling holes maybe disposed in the OD platform proximate the suction side. The firstgroup of the plurality of cooling holes may comprise four cooling holes.A second group of the plurality of cooling holes may be disposed in theOD platform proximate the pressure side and near a leading edge of theairfoil. A second group of the plurality of cooling holes may comprisefive cooling holes disposed within two rows. The two rows may beperpendicular to each other. A third group of the plurality of coolingholes may be disposed in the OD platform proximate the suction side andnear a center of the airfoil. The third group of the plurality ofcooling holes may comprise two cooling holes. A fourth group of theplurality of cooling holes may be disposed in the OD platform proximatethe suction side and near a trailing edge of the airfoil. The fourthgroup of the plurality of cooling holes may comprise two cooling holes.A fifth group of the plurality of cooling holes may be disposed in theOD platform proximate a leading edge of the airfoil. The fifth group ofthe plurality of cooling holes may comprise three cooling holes. Theplurality of cooling holes may be configured to eject coolant in adirection that film cools gaspath exposed surfaces of the OD platform.The plurality of cooling holes may be formed in substantial conformancewith platform cooling hole locations described by the set of Cartesiancoordinates set forth in Table 1.

According to various embodiments, a gas turbine engine comprises acompressor section, a combustor section and a turbine section includinga plurality of airfoils, wherein each airfoil projects from an ODplatform. The OD platform may comprise a gaspath and a non-gaspathsurface. These surfaces may couple together via a plurality of coolingholes. The plurality of cooling holes may comprise a plurality of groupsof cooling holes disposed in the OD platform proximate a suction side ofthe airfoil and a pressure side of the airfoil. The OD platform maycomprise five groups of cooling holes. Two of the five groups of coolingholes may be disposed in the OD platform on the pressure side andproximate a leading edge of the airfoil. Three of the five groups ofcooling holes are disposed in the OD platform proximate the suction sideof the airfoil. A first group of cooling holes disposed in the ODplatform on the suction side of the airfoil may be located proximate theleading edge of the airfoil. A second group of cooling holes disposed inthe OD platform on the suction side of the airfoil may be locatedtowards a middle section of the airfoil. A third group of cooling holesdisposed in the OD platform on the suction side of the airfoil may belocated proximate a trailing edge of the airfoil. The plurality ofcooling holes may be formed in substantial conformance with the ODplatform cooling hole locations described by the set of Cartesiancoordinates set forth in Table 1.

According to various embodiments, an apparatus, including an outerdiameter platform comprising: a non-gaspath face and a gaspath facecoupled to the non-gaspath face via a plurality of cooling holes isdisclosed. The apparatus may include an airfoil extending from the outerdiameter platform. The plurality of cooling holes are formed insubstantial conformance with platform cooling hole locations describedby the set of Cartesian coordinates set forth in Table 1.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates an example gas turbine engine.

FIG. 2 illustrates an OD platform with cooling holes in accordance withvarious embodiments.

FIG. 3 illustrates an OD platform having a cooling hole normal to asurface.

FIG. 4 illustrates an OD platform having a cooling hole angled from asurface.

FIG. 5 illustrates a perspective view of a turbine vane for a highpressure turbine section.

FIG. 6 illustrates a schematic view of a cooling hole layout inaccordance with various embodiments.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration. While these exemplary embodiments are described insufficient detail to enable those skilled in the art to practice theinventions, it should be understood that other embodiments may berealized and that logical changes and adaptations in design andconstruction may be made in accordance with this invention and theteachings herein. Thus, the detailed description herein is presented forpurposes of illustration only and not of limitation. The scope of theinvention is defined by the appended claims. For example, the stepsrecited in any of the method or process descriptions may be executed inany order and are not necessarily limited to the order presented.Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Also, any reference to attached, fixed, connected orthe like may include permanent, removable, temporary, partial, fulland/or any other possible attachment option. Additionally, any referenceto without contact (or similar phrases) may also include reduced contactor minimal contact. Surface shading lines may be used throughout thefigures to denote different parts but not necessarily to denote the sameor different materials. In some cases, reference coordinates may bespecific to each figure.

As used herein, “aft” refers to the direction associated with the tail(e.g., the back end) of an aircraft, or generally, to the direction ofexhaust of the gas turbine. As used herein, “forward” refers to thedirection associated with the nose (e.g., the front end) of an aircraft,or generally, to the direction of flight or motion.

In various embodiments and with reference to FIG. 1, a gas turbineengine 20 is provided. Gas turbine engine 20 may be a two-spool turbofanthat generally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mayinclude, for example, an augmenter section among other systems orfeatures. In operation, fan section 22 can drive air along a bypassflow-path B while compressor section 24 can drive air for compressionand communication into combustor section 26 then expansion throughturbine section 28. Although depicted as a turbofan gas turbine engine20 herein, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of gas turbine engines including three-spool architectures.

Gas turbine engine 20 may generally comprise a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A-A′ via one or more bearing systems 38 (shown asbearing system 38-1 and bearing system 38-2 in FIG. 2). It should beunderstood that various bearing systems 38 at various locations mayalternatively or additionally be provided, including for example,bearing system 38, bearing system 38-1, and bearing system 38-2.

Low speed spool 30 may generally comprise an inner shaft 40 thatinterconnects a fan 42, a low pressure (or first) compressor section 44(also referred to a low pressure compressor) and a low pressure (orfirst) turbine section 46. Inner shaft 40 may be connected to fan 42through a geared architecture 48 that can drive fan 42 at a lower speedthan low speed spool 30. Geared architecture 48 may comprise a gearassembly 60 enclosed within a gear housing 62. Gear assembly 60 couplesinner shaft 40 to a rotating fan structure. High speed spool 32 maycomprise an outer shaft 50 that interconnects a high pressure compressor52 (e.g., a second compressor section) and high pressure (or second)turbine section (“HPT”) 54. A combustor 56 may be located between highpressure compressor 52 and HPT 54. A mid-turbine frame 57 may be locatedgenerally between HPT 54 and low pressure turbine 46. Mid-turbine frame57 may support one or more bearing systems 38 in turbine section 28.Inner shaft 40 and outer shaft 50 may be concentric and rotate viabearing systems 38 about the engine central longitudinal axis A-A′,which is collinear with their longitudinal axes. As used herein, a “highpressure” compressor or turbine experiences a higher pressure than acorresponding “low pressure” compressor or turbine.

The core airflow may be compressed by low pressure compressor 44 thenhigh pressure compressor 52, mixed and burned with fuel in combustor 56,then expanded over HPT 54 and low pressure turbine 46. Mid-turbine frame57 includes airfoils 59 which are in the core airflow path. Low pressureturbine 46 and HPT 54 rotationally drive the respective low speed spool30 and high speed spool 32 in response to the expansion.

Gas turbine engine 20 may be, for example, a high-bypass geared aircraftengine. In various embodiments, the bypass ratio of gas turbine engine20 may be greater than about six (6). In various embodiments, the bypassratio of gas turbine engine 20 may be greater than ten (10). In variousembodiments, geared architecture 48 may be an epicyclic gear train, suchas a star gear system (sun gear in meshing engagement with a pluralityof star gears supported by a carrier and in meshing engagement with aring gear) or other gear system. Geared architecture 48 may have a gearreduction ratio of greater than about 2.3 and low pressure turbine 46may have a pressure ratio that is greater than about 5. In variousembodiments, the bypass ratio of gas turbine engine 20 is greater thanabout ten (10:1). In various embodiments, the diameter of fan 42 may besignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 may have a pressure ratio that is greaterthan about (5:1). Low pressure turbine 46 pressure ratio may be measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of low pressure turbine 46 prior to an exhaust nozzle. Itshould be understood, however, that the above parameters are exemplaryof various embodiments of a suitable geared architecture engine and thatthe present disclosure contemplates other gas turbine engines includingdirect drive turbofans.

Typically, vanes are cooled by communicating bleed air from thecompressor section through a bypass duct and into the OD platform. Thiscooling air is then directed through cooling hole features to providefilm cooling and internal convective cooling to reduce the vaneoperating metal temperature. Cooling hole placement design contributesto effective vane cooling. Accordingly, an efficient arrangement ofcooling holes assists in increasing cooling effectiveness.

In various embodiments, with reference to FIG. 2 and FIG. 3, and withcontinued reference to FIG. 1, an OD platform 202 is provided. FIG. 2schematically illustrates a proximal view of a HPT 54 vane OD platform202. The OD platform 202 is defined as having a forward face 292 and anaft face 294, as well as a non-gaspath face 297 and a gaspath face 296.The non-gaspath face 297 is located on the back side of the OD platform202 and is spaced apart opposite of the gaspath face 296. The airfoil206 extends inwardly from the OD platform 202.

An airfoil component, such as a vane, according to various embodiments,comprises an airfoil 206 that extends inwardly from the OD platform 202.The OD platform 202 includes at least one cooling hole connecting thegaspath face 296 and non-gaspath face 297 of the OD platform 202. Thecooling holes may couple with a plenum 295. As used herein, face andsurface, such as non-gaspath face and non-gaspath surface, may be usedinterchangeably.

In various embodiments and with reference to FIG. 2 and FIG. 3, the nextgeneration of turbofan engines may be designed for higher efficiencywhich is associated with higher pressure ratios and higher temperaturesin the HPT 54 (as shown in FIG. 1). Extreme operating temperatures andpressure ratios may create operating environments that may cause thermalloads that are higher than the thermal loads encountered in conventionalturbofan engines, in particular the forward face 292 of the OD platform202 located directly aft of the combustor section 26 (as shown in FIG.1). Hot exhaust air “E” flows from the combustor section 26 to theforward face 292. Due to thermal fatigue, the vane is made of materialssuch as nickel-based superalloys to withstand extreme conditions.Cooling holes (220, 240, 260, 280, 290) may be disposed in the ODplatform 202 configured to direct cooling air “D” from the non-gaspathside of the OD platform 202 to enter into the exhaust air flow “E” onthe gaspath side of the OD platform 202 due to a favorable pressuregradient between the gaspath side and non-gaspath side of the ODplatform 202, the non-gaspath side having greater pressure than thegaspath side. These cooling holes (220, 240, 260, 280, and 290) assistwith decreasing the OD platform 202 metal temperatures throughconvection and/or film cooling.

The location of the cooling holes (220, 240, 260, 280, and 290) may bestrategically placed. The strategic placement may include ejecting acoolant from a plurality of cooling holes (220, 240, 260, 280, and 290)in the OD platform 202 in a direction that provides film cooling to agaspath face 296 of an airfoil component, such as a stator vane. Coolingholes 220 are located on the pressure side 224 of the OD platform 202near the leading edge 228. According to various embodiments, the coolingholes 220 may have any desired shape and/or size. For instance, coolingholes may be round in shape with about a 0.017 inch (0.43 mm) diameter.Cooling holes 240 are located proximate the leading edge side 228 of theOD platform 202. Cooling holes 260 are located on the suction side 226of the OD platform 202 near the leading edge 228 of the airfoil 206.Cooling holes 280 are located on the suction side 226 of the OD platformnear the center/middle section of the airfoil 206. Cooling holes 290 arelocated on the suction side 226 of the airfoil near the trailing edge ofthe airfoil 206. The cooling holes (220, 240, 260, 280, 290) are notnecessarily all the same diameter and may vary in size, shape and flowdirection orientation.

According to various embodiments and with renewed reference to FIG. 2,cooling holes 220 may comprise cooling holes 210, 212, 214, 216, and218. Cooling holes 240 may comprise cooling holes 232, 234, and 236.Cooling holes 260 may comprise cooling holes 253, 254, 256, and 258.Cooling holes 280 may comprise cooling holes 272, and 274. Cooling holes290 may comprise cooling holes 286, and 288.

Cooling holes 220 may comprise a plurality of rows of holes, (e.g.,cooling holes 210 and 212 may comprise a first row and cooling holes214, 216 and 218 may comprise a second row. The first row and second rowof cooling holes 220 may each be non-parallel rows. For instance, thefirst row and second row of cooling holes 220 may be substantiallyperpendicular to each other. Cooling holes 240 may be located closer tothe leading edge 228 of the airfoil 206 as compared with cooling holes220. Cooling hole 258 may be positioned closer to the airfoil 206 thanthe position of adjacent cooling hole 256. The position of cooling holes290 may be closer to the trailing edge of the airfoil 206 as comparedwith the position of cooling holes 280. Cooling hole 274 may bepositioned closer to the airfoil 206 than the position of adjacentcooling hole 272. Cooling hole 288 may be positioned closer to theairfoil 206 than the position of adjacent cooling hole 286.

In various embodiments and with reference to FIG. 2 and FIG. 3, the sizeand shape of cooling holes (220, 240, 260, 280, and 290) may vary.Cooling holes (220, 240, 260, 280, and 290) can be any variation ofcircular, ovular, rectangular, triangular, or any other shape in crosssection. With reference to FIG. 3, cooling hole 304, which may be anycooling hole (220, 240, 260, 280, 290), can be situated normal to the ODplatform 302 gaspath face 396. Cooling air travels along path “D” in theaft direction through the bypass duct 308 and can enter into coolinghole 304 and enter into the gaspath “E” section to cool the airfoilcomponent, such as a vane. With reference to FIG. 4, cooling hole 404,which may be any cooling hole (220, 240, 260, 280, and 290), may besituated at an angle Φ to the OD platform 402 gaspath face 496. Theangle Φ at which cooling hole 404 is oriented may be any desired anglesuch as an angle varying between 18 and 40 degrees.

The method of creating the cooling holes (220, 240, 260, 280, and 290)can include any method of drilling, boring, or cutting as well as anyother method known to persons of ordinary skill in the art. According tovarious embodiments, cooling holes (220, 240, 260, 280, and 290) areformed via laser drilled holes.

In various embodiments and with reference to FIG. 5, the ID platform 574includes a radially inwardly extending tab 586 providing a tangentialon-board injector (TOBI) pin hole that is used to fasten the vane 560 toa TOBI 590. The center of the TOBI pin hole corresponds to a point P,which provides the reference zero-coordinate for the vane cooling holes.

In various embodiments the OD platform 202 includes cooling holes thatare shown in FIG. 6. The cooling holes each include a diameter and/orfootprint (where the holes break the surface of the OD platform 202) of0.010-0.035 inch (0.25-0.89 mm), in one example. Generally, the coolingholes are directed away from the airfoil 206 and from the leading edge292 toward the trailing edge 294. The angle of the cooling holesrelative to the OD platform 202 surface may be any desired angle such asbetween about 10° to 45°, in one example. The locations of the coolingholes are described in terms of Cartesian coordinates defined along x, yand z axes, which respectively correspond to the axial (x),circumferential (y) and radial (z) directions shown in the Figures. Thelocations of the OD platform 202 cooling holes correspond to thecenterline of holes where the holes break the surface of the platform.The coordinates of the cooling holes are respectively set forth in Table1 (in inches), which provide the nominal axial (x), circumferential (y)and radial (z) coordinates relative to the point P on a cold, uncoated,stationary vane. Each row in Table 1 corresponds to a cooling holelocation. The coordinates can be converted to metric (mm) by multiplyingby 25.4, or could be converted to any other units. A fillet is providedbetween the airfoil 206 and the OD platform 202. The half of the filletadjacent to each platform is generally considered to be a part of ODplatform 202.

TABLE 1 Hole No. X Y Z 212 0.406 0.808 2.493 214 0.278 0.884 2.504 2160.263 0.747 2.508 218 0.250 0.597 2.510 232 0.237 0.138 2.500 234 0.2220.292 2.509 236 0.237 0.450 2.510 253 1.326 0.440 2.347 254 1.187 0.3242.356 256 1.050 0.209 2.377 258 0.913 0.094 2.394 272 1.424 0.982 2.341274 1.274 0.875 2.344 286 1.431 1.217 2.327 288 1.313 1.250 2.321

The coordinates define break out points (illustrated in FIG. 6 as thecenter of small circles) of the cooling holes on a cold, uncoated,stationary vane. Additional elements such as additional cooling holes,protective coatings, fillets and seal structures may also be formed ontothe specified platform surface, or onto an adjacent airfoil surface, butthese elements are not necessarily described by the coordinates.

Due to manufacturing tolerances, the external breakout of the centerlineof the hole can fall within a 0.200 inch (5.08 mm) diameter circleinscribed on the surface of the part. The tolerances about the coolingholes 220, 240, 260, 280, 290 are schematically illustrated by thecircles circumscribing the cooling holes. However, the edge-to-edgespacing between adjacent cooling holes of at least 0.015 inch (0.381 mm)must be maintained. Additionally, the cooling holes are also locatedwithin the OD platform 202 regardless of the tolerance. These tolerancesare generally constant or not scalable, and apply to the specifiedplatform surface, regardless of size.

Substantial conformance is based on points representing the cooling holelocations, for example in inches or millimeters, as determined byselecting particular values of the scaling parameters. A substantiallyconforming airfoil, blade or vane structure has cooling holes thatconform to the specified sets of points, within the specified tolerance.

Alternatively, substantial conformance is based on a determination by anational or international regulatory body, for example in a partcertification or part manufacture approval (PMA) process for the FederalAviation Administration, the European Aviation Safety Agency, the CivilAviation Administration of China, the Japan Civil Aviation Bureau, orthe Russian Federal Agency for Air Transport. In these configurations,substantial conformance encompasses a determination that a particularpart or structure is identical to, or sufficiently similar to, thespecified airfoil, blade or vane, or that the part or structure issufficiently the same with respect to a part design in a type-certifiedor type-certificated airfoil, blade or vane, such that the part orstructure complies with airworthiness standards applicable to thespecified blade, vane or airfoil. In particular, substantial conformanceencompasses any regulatory determination that a particular part orstructure is sufficiently similar to, identical to, or the same as aspecified blade, vane or airfoil, such that certification orauthorization for use is based at least in part on the determination ofsimilarity.

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. Furthermore, theconnecting lines shown in the various figures contained herein areintended to represent exemplary functional relationships and/or physicalcouplings between the various elements. It should be noted that manyalternative or additional functional relationships or physicalconnections may be present in a practical system. However, the benefits,advantages, solutions to problems, and any elements that may cause anybenefit, advantage, or solution to occur or become more pronounced arenot to be construed as critical, required, or essential features orelements of the inventions. The scope of the inventions is accordinglyto be limited by nothing other than the appended claims, in whichreference to an element in the singular is not intended to mean “one andonly one” unless explicitly so stated, but rather “one or more.”Moreover, where a phrase similar to “at least one of A, B, or C” is usedin the claims, it is intended that the phrase be interpreted to meanthat A alone may be present in an embodiment, B alone may be present inan embodiment, C alone may be present in an embodiment, or that anycombination of the elements A, B and C may be present in a singleembodiment; for example, A and B, A and C, B and C, or A and B and C.Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “one embodiment”, “an embodiment”,“various embodiments”, etc., indicate that the embodiment described mayinclude a particular feature, structure, or characteristic, but everyembodiment may not necessarily include the particular feature,structure, or characteristic. Moreover, such phrases are not necessarilyreferring to the same embodiment. Further, when a particular feature,structure, or characteristic is described in connection with anembodiment, it is submitted that it is within the knowledge of oneskilled in the art to affect such feature, structure, or characteristicin connection with other embodiments whether or not explicitlydescribed. After reading the description, it will be apparent to oneskilled in the relevant art(s) how to implement the disclosure inalternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f), unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises”,“comprising”, or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

What is claimed is:
 1. An airfoil component comprising: an outerdiameter platform comprising: a non-gaspath face; a gaspath face coupledto the non-gaspath face via a plurality of cooling holes; and an airfoilextending from the outer diameter platform, wherein the plurality ofcooling holes comprise a plurality of groups of cooling holes disposedin the outer diameter platfortii proximate a suction side of the airfoiland a pressure side of the airfoil, wherein a plenum is formed adjacentto the non-gaspath face, wherein the plurality of cooling holes arecoupled to the plenum; wherein a second group of the plurality ofcooling holes is disposed in the outer diameter platform proximate thepressure side and near a leading edge of the airfoil, wherein the secondgroup of the plurality of cooling holes has five cooling holes disposedwithin two rows.
 2. The airfoil component of claim 1, wherein theairfoil component is a first stage turbine vane.
 3. The airfoilcomponent of claim 1, wherein a first group of the plurality of coolingholes is disposed in the outer diameter platform proximate the suctionside.
 4. The airfoil component of claim 3, wherein the first group ofthe plurality of cooling holes has four cooling holes.
 5. The airfoilcomponent of claim 1, wherein the two rows are perpendicular to eachother.
 6. The airfoil component of claim 1, wherein a third group of theplurality of cooling holes is disposed in the outer diameter platformproximate the suction side and near a center of the airfoil.
 7. Theairfoil component of claim 6, wherein the third group of the pluralityof cooling holes has two cooling holes.
 8. The airfoil component ofclaim 1, wherein a fourth group of the plurality of cooling holes isdisposed in the outer diameter platform proximate the suction side andnear a trailing edge of the airfoil.
 9. The airfoil component of claim8, wherein the fourth group of the plurality of cooling holes has twocooling holes.
 10. The airfoil component of claim 1, wherein a fifthgroup of the plurality of cooling holes is disposed in the outerdiameter platform proximate a leading edge of the airfoil.
 11. Theairfoil component of claim 10, wherein the fifth group of the pluralityof cooling holes has three cooling holes.
 12. The airfoil component ofclaim 1, wherein the plurality of cooling holes are configured to ejectcoolant in a direction that film cools gaspath exposed surfaces of theouter diameter platform.
 13. The airfoil component of claim 1, whereinthe plurality of cooling holes are formed in substantial conformancewith outer diameter platform cooling hole locations, wherein thelocations of the cooling holes are described in terms of Cartesiancoordinates defined by along x, y, and z axes and the coordinates of thecooling holes are: Hole No. X Y Z First Hole 0.406 0.808 2.493 SecondHole 0.278 0.884 2.504 Third Hole 0.263 0.747 2.508 Fourth Hole 0.2500.597 2.510 Fifth Hole 0.237 0.138 2.500 Sixth Hole 0.222 0.292 2.509Seventh Hole 0.237 0.450 2.510 Eighth Hole 1.326 0.440 2.347 Ninth Hole1.187 0.324 2.356 Tenth Hole 1.050 0.209 2.377 Eleventh Hole 0.913 0.0942.394 Twelfth Hole 1.424 0.982 2.341 Thirteenth Hole 1.274 0.875 2.344Fourteenth Hole 1.431 1.217 2.327 Fifteenth Hole 1.313 1.250  2.321.


14. A gas turbine engine comprising: a compressor section; a combustorsection; and a turbine section including a plurality of airfoils,wherein each airfoil projects from an outer diameter platform; the outerdiameter platform comprising: a non-gaspath face; and a gaspath facecoupled to the non-gaspath face via a plurality of cooling holes,wherein the plurality of cooling holes comprise a plurality of groups ofcooling holes disposed in the outer diameter platform proximate asuction side of a vane and a pressure side of the vane: wherein a secondgroup of the plurality of cooling holes is disposed in the outerdiameter platform proximate the pressure side and near a leading edge ofthe airfoil, wherein the second group of the plurality of cooling holeshas five cooling holes disposed within two rows.
 15. The gas turbineengine of claim 14, further comprising five groups of cooling holes. 16.The gas turbine engine of claim 15, wherein two of the five groups ofcooling holes are disposed in the outer diameter platform on thepressure side and proximate a leading edge of the airfoil, wherein threeof the five groups of cooling holes are disposed in the outer diameterplatform proximate the suction side of the airfoil, wherein a firstgroup of cooling holes disposed in the outer diameter platform on thesuction side of the airfoil is located proximate the leading edge of theairfoil, wherein a second group of cooling holes disposed in outerdiameter platform on the suction side of the airfoil is located towardsa middle section of the airfoil, and wherein a third group of coolingholes disposed in the outer diameter platform on the suction side of theairfoil are located proximate a trailing edge of the airfoil.
 17. Thegas turbine engine of claim 14, wherein the plurality of cooling holesare formed in substantial conformance with outer diameter platformcooling hole locations, wherein the locations of the cooling holes aredescribed in terms of Cartesian coordinates defined by along x, y, and zaxes and the coordinates of the cooling holes are: Hole No. X Y Z FirstHole 0.406 0.808 2.493 Second Hole 0.278 0.884 2.504 Third Hole 0.2630.747 2.508 Fourth Hole 0.250 0.597 2.510 Fifth Hole 0.237 0.138 2.500Sixth Hole 0.222 0.292 2.509 Seventh Hole 0.237 0.450 2.510 Eighth Hole1.326 0.440 2.347 Ninth Hole 1.187 0.324 2.356 Tenth Hole 1.050 0.2092.377 Eleventh Hole 0.913 0.094 2.394 Twelfth Hole 1.424 0.982 2.341Thirteenth Hole 1.274 0.875 2.344 Fourteenth Hole 1.431 1.217 2.327Fifteenth Hole 1.313 1.250  2.321.


18. An apparatus, comprising: an outer diameter platform comprising: anon-gaspath face; and a gaspath face coupled to the non-gaspath face viaa plurality of cooling holes; and an airfoil extending from the outerdiameter platform, wherein the plurality of cooling holes are formed insubstantial conformance with platform cooling hole locations, whereinthe locations of the cooling holes are described in terms of Cartesiancoordinates defined by along x, y, and z axes and the coordinates of thecooling holes are: Hole No. X Y Z First Hole 0.406 0.808 2.493 SecondHole 0.278 0.884 2.504 Third Hole 0.263 0.747 2.508 Fourth Hole 0.2500.597 2.510 Fifth Hole 0.237 0.138 2.500 Sixth Hole 0.222 0.292 2.509Seventh Hole 0.237 0.450 2.510 Eighth Hole 1.326 0.440 2.347 Ninth Hole1.187 0.324 2.356 Tenth Hole 1.050 0.209 2.377 Eleventh Hole 0.913 0.0942.394 Twelfth Hole 1.424 0.982 7.341 Thirteenth Hole 1.274 0.875 2.344Fourteenth Hole 1.431 1.217 2.327 Fifteenth Hole 1.313 1.250  2.321.


19. The apparatus of claim 18, wherein the plurality of cooling holescomprise a plurality of groups of cooling holes disposed in the outerdiameter platform proximate a suction side of a vane and a pressure sideof the vane; wherein a second group of the plurality of cooling holes isdisposed in the outer diameter platform proximate the pressure side andnear a leading edge of the airfoil, wherein the second group of theplurality of cooling holes has five cooling holes disposed within tworows.